• Journal of the Korean Society of Combustion
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• v.22 no.1
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• pp.46-52
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• 2017

High-frequency combustion instabilities may occur during the development of feasible engine combustors. These instabilities can result in irreparable damages to the wall of combustors or the degradation of engine performance. So, it is essential to identify injectors that have high stability characteristics during the early stages of development. The objective of present study was to assess the stability of coaxial injectors and an impinging injector with different recess lengths in order to develop stable injectors optimally. Stability margin was evaluated based on the distance from operating condition to the unstable regions. A simulating combustion test method was used to analyze the stability of injectors. A small-scale combustion chamber was designed to simulate the first tangential acoustic mode of the actual combustor. Gaseous oxygen and a mixture of methane and propane were used as simulant propellants to satisfy their flow similarity to the actual propellants of a combustor in a liquid rocket combustor. The results indicated that injectors having small recess lengths showed relatively large combustion stability margins. For the injectors of large recess lengths, instability regions with large and super-large amplitude oscillations were observed. Thus, injector with shorter recess lengths had a higher stability than that of longer one due to the different mixing processes.

• 한국연소학회:학술대회논문집
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• 2014.11a
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• pp.153-156
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• 2014

This study was aimed to assess combustion stability for coaxial swirl injector and FOOF impinging injector which would be candidates in liquid rocket engine combustors. Simulating combustion tests under atmospheric condition have been conducted by gaseous oxygen and the mixture of methane and propane, using two actual injectors. By analyzing the measured dynamic pressure signals, we have evaluated the combustion stability margin of both injectors by drawing a stability map.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.207-210
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• 2005

The results of combustion performance test of fuel-rich gas generator with dual swirl injectors are described. By changing simulating duct and recess number(RN) of the injectors, we inspected whether the combustion instability took place. When the injectors of RN = 0.5 were used, combustion instabilities could be reduced using the simulating duct. However, the effect of the simulating duct on the gas generator with the injectors of RN = 1.5 was not significant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.31-34
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• 2007

Performance tests of supersonic exhaust diffuser were conducted by using hot combustion gas for simulating high-altitude environment. The test rig consists of a combustion chamber, a vacuum chamber, water cooling ring and diffuser. Before combustion experiments, the preliminary leak tests were carried out on the liquid rocket engine and diffuser by using high pressure nitrogen(30barg) and a vacuum pump. The leak test results showed that there was no leaks at high pressure and vacuum pressure conditions.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.12
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• pp.1195-1201
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• 2010

Performance tests of a supersonic exhaust diffuser were conducted by using a liquid rocket engine for simulating high-altitude environment. The experimental setup consisted of a combustion chamber, a vacuum chamber and a diffuser. The combustion tests for simulating high-altitude environment were carried out at three cases by chamber pressure variation(26, 29, 32barg). The test results showed that the diffuser was started at all case and vacuum chamber pressures were approximately 140torr. The starting pressure using combustion gas was similar with that of cold gas, but the vacuum chamber pressure was relatively high because of high temperature in the vacuum chamber. The results of this test can be used as an essential database for the design of real-scale high-altitude simulation test facility in the future.

• 한국연소학회:학술대회논문집
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• 2004.11a
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• pp.141-145
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• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess length have been tested for the comparative study of CH flame structure and dynamic characteristics. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a full-scale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.4
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• pp.81-88
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• 2013

It was designed and tested at the design point that an oxidizer rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burned with kerosene and the rest of LOx injected from injection holes in the regenerative cooling chamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 MPa. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidizer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.245-250
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• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess number have been tested for the comparative study of stability characteristic and flame structure. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a fullscale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.3
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• pp.1-7
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• 2018

A combustion test for a supersonic combustor was conducted using a direct-connected type supersonic combustor test facility. The facility was verified for the capability of simulating required flow conditions. The test condition was maintained at Mach 2.0, $915^{\circ}C$ and 496 kPa for 15 s. Using gaseous hydrogen as the fuel, the combustor model was also tested for its ignition and flame holding capability at the fuel equivalence ratio of 0.12. Combustion efficiency was 71%, and the supersonic flow regime was obtained at this test condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.73-80
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• 2018

At the early stages of development of high-speed propulsion systems, associated uncertainties cannot be easily modeled into probabilistic distributions, owing to the lack of test data, cost, and difficulty of simulating real-flight environments on the ground. To tackle this issue, in this research, the combustion efficiencies of dual-combustion ramjet engines are assumed to have been provided by experts and quantified by evidence theory. Using quantified uncertainty, the inlet area and combustor exit are optimized while satisfying reliability margins of thrust and thermal choking. The result shows a reasonable design of the engine under uncertain circumstances.